Composite structures are now extensively utilized in many industries and in some cases provide higher strength and stiffness than metallic materials. Composites also have a strength to weight ratio desirable in many aerospace applications. In addition, composites have a high resistance to most chemical and environmental threats.
Composites are formed with a variety of techniques most of which involve different ways of forming a combination fiber matrix material. One such technique is the resin transfer molding process: dry fiber material is formed to a desired part shape (preformed) and then impregnated with resin which is subsequently cured. Alternatively, a sizer or tackifier may be used to retain the dry fiber "preform" in the desired shape and the preform is subsequently impregnated with resin and cured similar to the procedure in the resin transfer molding process. Another technique uses prepreg materials in conjunction with a secondary molding operation. Fiber material is precombined with resin as a prepreg which is then shaped and cured in an autoclave, platen press, or suitable processing equipment which subjects the prepreg to various pressure and temperature cycles.
Also, composites can be combined during manufacture with materials other than fiber and resin as desired for added strength and flexibility. For example, foam layers or other core material can be incorporated between fiber and resin layers to form a cored or sandwich structure. This type of structure is usually applied when added stiffness is more critical than weight. The core is lighter than the composite providing added stiffness at reduced weight.
Regardless of the process used, however, all composites exhibit certain shortcomings. For example, interlaminar strength is poor in comparison with in plane properties of the composite because they are matrix dominated. Composites have excellent properties along the X-Y plane of the reinforcing fibers and manufacturing methods used to produce two dimensional structures are well developed. Low interlaminar (perpendicular to the plane in the Z-direction) strength and fracture toughness, however, limits the use of composites in many applications. Also, cracks or delaminations caused by thermal effects, impact events, or the presence of holes or free edges in the composite may seriously reduce compressive and flexural loading capacities or cause delamination which may result in premature structural failure.
There are a number of techniques for overcoming these limitations in composites. The two most frequently used are toughened matrices and through the thickness reinforcing fibers. Toughened matrices, however, are often 1.5 to 2 times more expensive than baseline systems, have poor hot wet properties, and still may not prove to offer enough toughness for a successful part design. Several techniques have been developed for placing fibers through the thickness of composites to improve interlaminar properties.
Also, stitching, stapling, and needling techniques are known, but these methods may cause a significant reduction of in-plane properties, are difficult to implement within complex-shaped laminates, and limit the type of fiber which can be utilized for reinforcement. Stitching uses needles which are often in excess of 0.2" in width. When penetrating a fiber laminate with a needle of this size, significant cutting or damage is caused to the load carrying in plane fibers. This can create strength reductions in excess of 20%. In addition to the needle damage, stitching uses a continuous thread. The loop in the thread traversing from one stitch to the next "kinks" the inplane fibers of the top few plies creating significant strength loss. Because of the demanding bend radii of stitching, the fiber material that can be practically used are limited to glass and KEVLAR. These materials are not the most effective through thickness reinforcements for all applications, and moreover, KEVLAR has been known to absorb moisture.
Another recent technique for overcoming these problems and disadvantages is shown in U.S. Pat. No. 4,808,461. A plurality of reinforcing elements are disposed wholly in a direction perpendicular to the plane of a body of thermally decomposable material. This structure is then placed on a prepreg and subjected to an elevated temperature which decomposes the thermally decomposable material. Pressure is used to drive the reinforcing elements into the prepreg which is then cured. The final composite part will contain the perpendicularly disposed reinforcing elements which add strength at desired locations of the composite part. This technique has been called "Z-direction reinforcement."
Another method, directed at prepreg techniques, uses a pin carrier to drive pins perpendicularly into the prepreg which is then shaped and cured. These techniques, too, have shortcomings.
For example, reinforcing fibers introduced purely perpendicular to the plane of the composite significantly reduce the tendency for the laminate to peel apart (mode I fracture), but they do not as significantly improve on the possibility of shear or mode II dominated failures. This is because these loads generally occur parallel to the in plane fibers of the composite and the reinforcing fibers are normal to the inplane fibers.
Manually inserting reinforcing rods at an angle to the inplane fibers in laminates along critically high stress planes before the layup of the graphite-epoxy plies is cured is known, but the tedious insertion of each reinforcing rod and the experimentation and analysis required to predetermine the critically high stress planes make this technique labor intensive and hence costly. It is also known to manually drive individual thermosetting resin-impregnated fibrous reinforcements at an angle into a fibrous material lay-up by mechanical impact or similar tools. This technique also requires labor intensive and close tolerance manufacturing techniques not suitable for all applications.